A helicopter is a highly complex dynamic machine. As a consequence a major design consideration is the stresses on blades resulting from air forces during blade revolutions in forward flight, as well as Coriolis forces resulting from dynamic coupling of flapping and lead/lag motion which act on each rotor blade throughout a revolution. Spinning rotor blades are so severely stressed that early on, materials could not be found which could tolerate those stresses. The rotating blades go through a variety of cyclic position changes due to external vibratory forces. There are also other vibratory forces associated with the rotor shaft motion when the helicopter is sitting on the ground and naturally rocking on its landing gear. In fact, these vibrations can become so severe they cannot be tolerated.
To some degree conventional articulated helicopter rotor hubs solve the vibration problem. They permit flapping motion of the blades, pitching movement of the blades, and lead/lag movement of the blades which relieves the vibratory bending moments into the hub, but this is accomplished at the expense of high hub weight since the hinges require lubrication and constant observation to detect wear and fatigue.
Hingeless helicopter rotor assemblies, generally employing composite materials, were developed to solve the weight and maintenance problems. They provide light weight helicopter hub assemblies. Such hingeless helicopter rotor assemblies usually include a rigid central hub member and radial flexible beams (flexbeams) rigidly attached by their root (inboard) ends to the hub. The blades are rigidly attached at their root ends to the outboard ends of the flexbeams.
Flexbeams can be in the form of C-beams, I-beams, T-beams, X-beams, and the like and one or more can be employed. A desirable flexbeam embodiment comprises two back-to-back C-beam members. The flexbeam is designed to bend in the vertical mode to accomodate blade flapping and in the horizontal mode to accomodate lead/lag motion of the blades. A rigid pitch shaft disposed between the two C-beam members transmits blade pitch-changing inputs from a control rod near the hub to the root end of the blade, and the flexbeams yield torsionally to permit the blade pitch changes. Flexbeams, thus, must be resilient in three orthogonal axes (pitch, lead/lag and flap). They are, of course, longitudinally rigid so that, by gimbal or other means, they transmit a blade centrifugal force to the central hub. Such a hingeless helicopter rotor with an elastic gimbal hub is illustrated in U.S. Pat. No. 4,323,332.
The use of hinged and flexbeam rotors does not completely eliminate the dynamic instability problem. Even using these designs energy must be absorbed. If damping means are not included, a self-excited vibration can cause the helicopter to vibrate with increasing amplitude, and the blades to move back and forth in the plane of rotation, but out of pattern. This phenomenon, known as ground resonance because it happens on the ground, must be prevented by damping.
The selection of the helicopter rotor damping means depends upon whether the rotor assembly is a hinge or hingeless type. A damping means for a hinge-type helicopter is exemplified in U.S. Pat. No. 2,554,774. Flexbeam rotor damping is described in U.S. Pat. No. 4,645,423. In the damping means described in that patent, damper assemblies are employed which include inboard and outboard housing units containing elastomeric damping layers. The damping layers deform in the shear mode, damping out the translatory input to the damper assembly. By the practice of the current invention, the shear principle is utilized without the weight and complexity of the linkages, housing units and long damper rods.